Turbine blade airfoil profile

ABSTRACT

A turbine blade for a gas turbine engine has an airfoil including leading and trailing edges joined by spaced-apart pressure and suction sides to provide an external airfoil surface extending from a platform in a spanwise direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil defined by a set of Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location.

TECHNICAL FIELD

The disclosure relates generally to gas turbine engines and, moreparticularly, to an airfoil that may be incorporated into a gas turbineengine.

BACKGROUND OF THE ART

Every compressor and turbine stage of a gas turbine engine must meet aplurality of design criteria to assure the best possible overall engineefficiency. The design goals dictate specific thermal and mechanicalrequirements that must be met pertaining to heat loading, parts life andmanufacturing, use of combustion gases, throat area, vectoring, theinteraction between stages to name a few. The design criteria for eachstage is constantly being re-evaluated and improved upon. Each airfoilis subject to flow regimes which lend themselves easily to flowseparation, which tend to limit the amount of work transferred to thecompressor, and hence the total power capability of the engine.Therefore, improvements in airfoil design are sought.

SUMMARY

In one aspect, there is provided a turbine blade for a gas turbineengine, the turbine blade comprising an airfoil including a leading anda trailing edge joined by a pressure and a suction side to provide anexternal airfoil surface extending from a platform in a spanwisedirection to a tip. The external airfoil surface is formed insubstantial conformance with multiple cross-sectional profiles of theairfoil defined by a set of Cartesian coordinates set forth in Table 1,the Cartesian coordinates provided by an axial coordinate scaled by alocal axial chord, a circumferential coordinate scaled by the localaxial chord, and a span location, wherein the local axial chordcorresponds to a width of the airfoil between the leading and trailingedges at the span location.

In another aspect, there is provided a gas turbine engine comprising alow pressure turbine. The low pressure turbine is configured to drive alow pressure compressor. The low pressure turbine comprises at least onestage of turbine blades, wherein at least one of the turbine blades ofthe at least one stage comprises an airfoil having leading and trailingedges joined by spaced-apart pressure and suction sides to provide anexternal airfoil surface extending from a platform in a span directionto a tip. The external airfoil surface is formed in substantialconformance with multiple cross-section profiles of the airfoil definedby a set of Cartesian coordinates set forth in Table 1. The Cartesiancoordinates are provided by an axial coordinate scaled by a local axialchord, a circumferential coordinate scaled by the local axial chord, anda span location, wherein the local axial chord corresponds to a width ofthe airfoil between the leading and trailing edges at the span location.

In a further aspect, there is provided a low pressure turbine bladecomprising: a platform and an airfoil extending in a spanwise directionfrom the platform to a tip. The airfoil has an external airfoil surfaceformed in substantial conformance with multiple cross-section airfoilprofiles defined by a set of Cartesian coordinates set forth in Table 1.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic cross-section view of a gas turbine engine;

FIG. 2 is a cross-section side view of a low pressure (LP) turbinesection of the gas turbine engine shown in FIG. 1 ;

FIG. 3 a is an isometric view of a shrouded LP turbine blade;

FIG. 3 b is a plan, top view of the turbine blade of FIG. 3 aillustrating directional references, the tip shroud of the blade omittedfor clarity;

FIG. 4 a is an isometric view of an exemplary turbine airfoilcorresponding to the directional references of FIG. 3 b;

FIG. 4 b is a pressure side view of the exemplary turbine airfoil shownin FIG. 4 a;

FIG. 4 c is a suction side view of the exemplary turbine airfoil shownin FIG. 4 a ; and

FIG. 5 depicts the span positions and local axial chords reference inTable 1.

DETAILED DESCRIPTION

FIG. 1 illustrates an example of a gas turbine engine 10 of a typepreferably provided for use in subsonic flight, generally comprising inserial flow communication an air inlet 11, a compressor 12 forpressurizing the air from the air inlet 11, a combustor 13 in which thecompressed air is mixed with fuel and ignited for generating an annularstream of hot combustion gases, a turbine 14 for extracting energy fromthe combustion gases, and an exhaust 15 through which the combustiongases exit the engine 10. According to the illustrated exemplary engine,the turbine 14 is drivingly connected to an input end of a reductiongearbox RGB 16. The RGB 16 has an output end drivingly connected to anoutput shaft 18 configured to drive a rotatable load (not shown). Therotatable load can, for instance, take the form of a propeller or arotor, such as a helicopter main rotor. Still according to theillustrated embodiment, all the compressor and the turbine rotors aremounted in-line for rotation about the engine centerline X. However, itis understood that the turbine and compressor rotors could havedifferent rotation axes. Also, it is understood that the conceptsdescribed herein are not limited to use with turboprop or turboshaftengines as the teachings may be applied to other types of turbineengines including various engine architectures. Furthermore, while theillustrated exemplary engine is a two-spool engine, it is understoodthat the engine could include a different number of spools. Forinstance, the engine could have 1 to 3 spools.

The exemplified engine 10 has an axially extending central core whichdefines an annular gaspath 20 through which gases flow, as depicted byflow arrows in FIG. 1 . The exemplary embodiment shown in FIG. 1 is a“through flow” engine because gases flow through the gaspath 20 from theair inlet 11 at a front portion of the engine 10, to the exhaust 15 at arear portion thereof. However, it is understood that the engine 10 couldadopt different configurations, including a reverse flow configuration,the engine configuration illustrated in FIG. 1 being provided forillustrative purposes only.

The terms “upstream” and “downstream” used herein refer to the directionof an air/gas flow passing through the gaspath 20 of the engine 10. Itshould also be noted that the terms “axial”, “radial”, “angular” and“circumferential” are used with respect to the rotation axes of theturbine and compressor rotors (i.e. the engine centerline X in theexemplary engine).

According to the illustrated embodiment, the turbine 14 comprises a lowpressure (LP) turbine 14 a and a high pressure (HP) turbine 14 b. The HPturbine 14 b is drivingly connected to an HP compressor 12 b via an HPshaft 22 b. The HP turbine 14 b, the HP shaft 22 b and the HP compressor12 b form one of the two spools of the engine 10, namely the HP spool.According to the illustrated embodiment, the HP turbine 14 b and the HPcompressor 12 b each have a single stage of rotating blades. However, itis understood that the HP turbine 14 b and the HP compressor 12 b couldhave any suitable number of stages.

Still according to the illustrated embodiment, the LP turbine 14 a isdrivingly connected to an LP compressor 12 a via an LP shaft 24 a. TheLP turbine 14 a, the LP shaft 24 a and the LP compressor 12 a form theother one of the two spools of the engine 10, namely the LP spool. TheHP spool and the LP spool are independently rotatable. According to theillustrated embodiment, the LP turbine 14 a has three stages of turbineblades, whereas the LP compressor 12 a has a single stage of LPcompressor blades. However, it is understood that the LP turbine 14 aand the LP compressor 12 a could have any suitable number of stages. Forinstance, according to one embodiment, the LP turbine 14 a is a twostage LP turbine.

The LP shaft 24 a is drivingly connected to the input end of the RBG 16to drive the output shaft 18. Accordingly, the LP turbine 14 a (alsoknown as the power turbine) can be used to drive both the LP compressor12 a and the output shaft 18. An additional gearbox or the like (notshown) can be provided between the LP compressor 12 a and the LP turbine14 a to allow the LP compressor 12 a to rotate at a different speed fromthe LP turbine 14 a.

In use, the air flowing through the inlet 11 is compressed by the LPcompressor 12 a then the HP compressor 12 b, mixed and burned with fuelin the combustor 13, then expanded over the HP turbine 14 b and the LPturbine 14 a before being discharged through the exhaust 15. The HPturbine 14 b drives the HP compressor 12 b, whereas the LP turbine 14 adrives the LP compressor 12 a and the output shaft 18.

Referring to FIG. 2 , it can be appreciated that the LP turbine 14 acomprises series of rotating blades and stationary vanes that extendsinto the gaspath 20 of the engine 10. In the exemplary LP turbine 14 a,first, second and third arrays of circumferentially spaced-apartstationary vanes are axially spaced-apart from one another along theaxis X. A first stage array 26 a of circumferentially spaced-apart LPturbine blades 28 a is disposed axially between the first and secondarrays 30 a, 32 a of turbine vanes 34 a, 36 a. A second stage array 38 aof circumferentially spaced-apart LP turbine blades 40 a is disposedaxially between the second and third arrays 32 a, 42 a of turbine vanes36 a, 44 a. A third stage array 46 a of circumferentially spaced-apartLP turbine blades 48 a is disposed downstream of the third array 42 a ofstationary turbine vanes 44 a. The second stage LP turbine blades 40 aare mounted in position along a stacking line corresponding to axis Z inFIG. 2 . The stacking line defines the axial location (x) where theblades 40 a are mounted along the centerline of the engine 10.

FIGS. 3 a and 3 b schematically illustrate an example of a shrouded LPturbine blade suitable for use as a first, second, third, fourth orfifth stage LP turbine blade. According to the illustrated example, theLP turbine blade is a second stage LP turbine blade of a 3-stage LPturbine. It can be seen that the second stage LP turbine blade 40 aincludes an airfoil 50 having an exterior/external surface 52 extendingin a chordwise direction CW between a leading edge 54 and a trailingedge 56 and in a spanwise direction Z from a platform 58 to a shroudedtip 60. The airfoil 50 is provided between pressure and suction sides62, 64 in an airfoil thickness direction T, which is generallyperpendicular to the chordwise direction CW. A root 66 depends radiallyinwardly from the platform 58 for detachably mounting the blade 40 a toa rotor disk. It is however understood that the turbine blade 40 a couldbe integrally formed with the disk. In such a configuration, the root iseliminated and the platform is provided at the outer diameter of therotor disk.

The external surface 52 of the airfoil 50 generates lift based upon itsgeometry and direct flow along the gaspath 20. Various views of theairfoil of the second stage low pressure turbine blade 40 a are shown inFIGS. 4 a-4 c . In one example, the second-stage array 38 a consists ofseventy-one (71) turbine blades 40 a, but the number may vary accordingto engine size. The turbine blades can be constructed from ahigh-strength, heat-resistant material such as a nickel-based orcobalt-based superalloy, or of a high-temperature, stress-resistantceramic or composite material, for example. In cooled configurations,internal fluid passages and external cooling apertures provide for acombination of impingement and film cooling. In addition, one or morethermal barrier coatings (TBC), abrasion-resistant coatings, and/orother protective coatings may be applied to the turbine blades.

Referring to FIG. 5 , the geometries of external surfaces of airfoil aredefined in terms of Cartesian coordinates defined along x, y, and zaxes, which respectively correspond to the axial (x), circumferential(y), and span/radial (Z-span) (z) directions shown in FIGS. 3 a and 3 b. The span coordinate is provided as a radial distance (ΔZI-ΔZ3) fromthe rotation axis X of the airfoil 50. The “0” span is taken at a pointP where the airfoil 50 meets the platform 58, as schematicallyillustrated in FIG. 5 . The overall or full span is 100% the distancefrom the point P to the tip 60 in the span direction Z-span. By way ofexample, the “¼ span” (ΔZ1) is 25% the distance from the point P towardthe tip 60 in the span direction Z-span.

The axial (x) and circumferential (y) coordinates are normalized by thelocal axial chord (Bx) for the (3) given span locations (ΔZ1-ΔZ3) (Thelocal axial chord (Bx1) for axial (x) and circumferential (y)coordinates associated with the ¼ span (ΔZI) corresponds to the width ofthe airfoil 50 between the leading and trailing edges 54, 56 at the ¼span location (ΔZI).

The contour of the airfoil is set forth in Table 1, which provides theaxial (x) and circumferential (y) coordinates (in inches) scaled by thelocal axial chord (Bx) for the given span locations or positions ΔZ1,ΔZ2 and ΔZ3 shown in FIG. 5 . 3-D airfoil surfaces are formed by joiningadjacent points in Table 1 in a smooth manner and joining adjacentsections or sectional profiles along the span. The manufacturingtolerance relative to the specified coordinates is ±0.050 inches (±1.27mm). The coordinates define points on a cold, uncoated, stationaryairfoil surface at nominal definition, in a plane at the correspondingspan positions. Additional elements such as cooling holes, protectivecoatings, fillets, and seal structures may also be formed onto thespecified airfoil surface, or onto an adjacent platform surface, butthese elements are not necessarily defined by the normalizedcoordinates. For example, a variable coating C may be applied between0.0001 inches (0.003 mm) (trace) and 0.01 inches (0.28 mm) thick.According to one particular embodiment, a constant coating of 0.0015inches (0.0381 mm) is applied.

TABLE 1 REFERENCE RADIUS: Z1 SECTION COORDINATES (X,Y)/BX1 0.000 0.0440.002 0.049 0.005 0.056 0.012 0.067 0.021 0.076 0.032 0.085 0.043 0.0940.055 0.104 0.075 0.117 0.100 0.132 0.125 0.145 0.150 0.155 0.175 0.1630.200 0.169 0.225 0.174 0.250 0.177 0.275 0.179 0.300 0.178 0.325 0.1770.350 0.174 0.375 0.170 0.400 0.164 0.425 0.157 0.450 0.149 0.475 0.1390.500 0.128 0.525 0.116 0.550 0.102 0.575 0.087 0.600 0.071 0.625 0.0530.650 0.034 0.675 0.014 0.700 −0.008   0.725 −0.031   0.750 −0.055  0.775 −0.081   0.800 −0.107   0.825 −0.134   0.850 −0.163   0.875−0.193   0.900 −0.223   0.925 −0.255   0.943 −0.279   0.959 −0.300  0.972 −0.318   0.982 −0.331   0.988 −0.340   0.993 −0.347   0.997−0.353   0.999 −0.357   1.000 −0.360   0.999 −0.367   0.997 −0.371  0.993 −0.374   0.988 −0.375   0.982 −0.374   0.972 −0.364   0.959−0.348   0.943 −0.328   0.925 −0.306   0.900 −0.275   0.875 −0.246  0.850 −0.217   0.825 −0.190   0.800 −0.165   0.775 −0.141   0.750−0.119   0.725 −0.098   0.700 −0.078   0.675 −0.060   0.650 −0.043  0.625 −0.027   0.600 −0.012   0.575 0.002 0.550 0.015 0.525 0.027 0.5000.038 0.475 0.048 0.450 0.057 0.425 0.065 0.400 0.072 0.375 0.078 0.3500.082 0.325 0.086 0.300 0.089 0.275 0.091 0.250 0.091 0.225 0.091 0.2000.089 0.175 0.085 0.150 0.080 0.125 0.073 0.100 0.065 0.075 0.054 0.0550.044 0.043 0.037 0.032 0.031 0.021 0.028 0.012 0.027 0.005 0.029 0.0020.034 REFERENCE RADIUS: Z2 SECTION COORDINATES (X,Y)/BX2 0.000 0.1370.002 0.148 0.005 0.156 0.012 0.166 0.021 0.176 0.032 0.184 0.043 0.1920.055 0.202 0.075 0.213 0.100 0.226 0.125 0.236 0.150 0.244 0.175 0.2490.200 0.252 0.225 0.254 0.250 0.253 0.275 0.251 0.300 0.247 0.325 0.2410.350 0.234 0.375 0.225 0.400 0.215 0.425 0.203 0.450 0.190 0.475 0.1760.500 0.159 0.525 0.142 0.550 0.123 0.575 0.103 0.600 0.081 0.625 0.0580.650 0.034 0.675 0.008 0.700 −0.019   0.725 −0.047   0.750 −0.077  0.775 −0.107   0.800 −0.139   0.825 −0.172   0.850 −0.205   0.875−0.240   0.900 −0.275   0.925 −0.312   0.943 −0.339   0.959 −0.362  0.972 −0.382   0.982 −0.397   0.988 −0.407   0.993 −0.415   0.997−0.421   0.999 −0.426   1.000 −0.430   0.999 −0.436   0.997 −0.440  0.993 −0.443   0.988 −0.445   0.982 −0.444   0.972 −0.436   0.959−0.418   0.943 −0.397   0.925 −0.372   0.900 −0.337   0.875 −0.304  0.850 −0.270   0.825 −0.238   0.800 −0.206   0.775 −0.176   0.750−0.147   0.725 −0.119   0.700 −0.092   0.675 −0.066   0.650 −0.042  0.625 −0.019   0.600 0.002 0.575 0.023 0.550 0.042 0.525 0.060 0.5000.076 0.475 0.092 0.450 0.106 0.425 0.119 0.400 0.130 0.375 0.140 0.3500.149 0.325 0.157 0.300 0.163 0.275 0.168 0.250 0.172 0.225 0.174 0.2000.174 0.175 0.173 0.150 0.170 0.125 0.165 0.100 0.158 0.075 0.149 0.0550.140 0.043 0.133 0.032 0.129 0.021 0.127 0.012 0.127 0.005 0.129 0.0020.134 REFERENCE RADIUS: Z3 SECTION COORDINATES (X,Y)/BX3 0.000 0.2570.002 0.267 0.005 0.274 0.012 0.284 0.021 0.293 0.032 0.301 0.043 0.3070.055 0.314 0.075 0.323 0.100 0.333 0.125 0.339 0.150 0.343 0.175 0.3450.200 0.345 0.225 0.343 0.250 0.339 0.275 0.333 0.300 0.325 0.325 0.3150.350 0.304 0.375 0.291 0.400 0.276 0.425 0.259 0.450 0.241 0.475 0.2210.500 0.200 0.525 0.177 0.550 0.153 0.575 0.126 0.600 0.099 0.625 0.0700.650 0.039 0.675 0.008 0.700 −0.025   0.725 −0.059   0.750 −0.093  0.775 −0.129   0.800 −0.166   0.825 −0.204   0.850 −0.243   0.875−0.282   0.900 −0.323   0.925 −0.365   0.943 −0.396   0.959 −0.423  0.972 −0.446   0.982 −0.464   0.988 −0.475   0.993 −0.485   0.997−0.492   0.999 −0.497   1.000 −0.503   0.999 −0.508   0.997 −0.512  0.993 −0.516   0.988 −0.518   0.982 −0.518   0.972 −0.512   0.959−0.491   0.943 −0.467   0.925 −0.439   0.900 −0.400   0.875 −0.362  0.850 −0.323   0.825 −0.285   0.800 −0.247   0.775 −0.209   0.750−0.173   0.725 −0.138   0.700 −0.103   0.675 −0.071   0.650 −0.039  0.625 −0.010   0.600 0.018 0.575 0.045 0.550 0.070 0.525 0.093 0.5000.115 0.475 0.135 0.450 0.154 0.425 0.171 0.400 0.187 0.375 0.202 0.3500.215 0.325 0.227 0.300 0.237 0.275 0.246 0.250 0.254 0.225 0.260 0.2000.264 0.175 0.267 0.150 0.269 0.125 0.267 0.100 0.264 0.075 0.258 0.0550.252 0.043 0.248 0.032 0.245 0.021 0.245 0.012 0.246 0.005 0.249 0.0020.255

This set of points represents a novel and unique solution to the targetdesign criteria mentioned herein above, and are well-adapted for use inthe first, second, third, fourth or fifth stage of an LP turbine. Forinstance, the blade airfoil defined by the coordinates in table 1 couldbe used in the second LP pressure turbine blade array 38 a ofexemplified engine 10. According to at least some embodiments, theturbine airfoil profile is particularly configured to improve theservice life of the second stage LP turbine blades 40 a.

In general, the turbine blade airfoil, as described herein, has acombination of axial sweep and tangential lean. Depending onconfiguration, the lean and sweep angles sometimes vary by up to ±10° ormore. In addition, the turbine blade is sometimes rotated with respectto a radial axis or a normal to the platform or shroud surface, forexample, by up to ±10° or more.

Novel aspects of the turbine blade and associated airfoil surfacesdescribed herein are achieved by substantial conformance to specifiedgeometries. Substantial conformance generally includes or may include amanufacturing tolerance of ±0.05 inches (±1.27 mm), in order to accountfor variations in molding, cutting, shaping, surface finishing and othermanufacturing processes, and to accommodate variability in coatingthicknesses. This tolerance is generally constant or not scalable, andapplies to each of the specified blade surfaces, regardless of size.

Substantial conformance is based on sets of points representing athree-dimensional surface with particular physical dimensions, forexample, in inches or millimeters, as determined by selecting particularvalues of the scaling parameters. A substantially conforming airfoil,blade or, or vane structure has surfaces that conform to the specifiedsets of points, within the specified tolerance.

Alternatively, substantial conformance is based on a determination by anational or international regulatory body, for example, in a partcertification or part manufacture approval (PMA) process for the FederalAviation Administration, the European Aviation Safety Agency, the CivilAviation Administration of China, the Japan Civil Aviation Bureau, orthe Russian Federal Agency for Air Transport. In these configurations,substantial conformance encompasses a determination that a particularpart or structure is identical to, or sufficiently similar to, thespecified airfoil, blade, or vane, or that the part or structurecomplies with airworthiness standards applicable to the specified blade,vane, or airfoil. In particular, substantial conformance encompasses anyregulatory determination that a particular part or structure issufficiently similar to, identical to, or the same as a specified blade,vane, or airfoil, such that certification or authorization for use isbased at least in part on the determination of similarity.

The embodiments described in this document provide non-limiting examplesof possible implementations of the present technology. Upon review ofthe present disclosure, a person of ordinary skill in the art willrecognize that changes may be made to the embodiments described hereinwithout departing from the scope of the present technology. For example,the present disclosure includes several aspects and embodiments thatinclude particular features. Although these particular features may bedescribed individually, it is within the scope of the present disclosurethat some or all of these features may be combined with any one of theaspects and remain within the scope of the present disclosure. Yetfurther modifications could be implemented by a person of ordinary skillin the art in view of the present disclosure, which modifications wouldbe within the scope of the present technology.

The invention claimed is:
 1. A turbine blade for a gas turbine engine,the turbine blade comprising: an airfoil having a leading and a trailingedge joined by a pressure and a suction side extending from a platformin a spanwise direction to a tip; and wherein the airfoil has a firstcross-section profile, a second cross-section profile and a thirdcross-section profile respectively at a first span location, a secondspan location and a third span location of the airfoil, the first,second and third cross-section profiles defined by an entirety ofCartesian coordinates set forth in Table 1, the Cartesian coordinatesprovided by an axial coordinate scaled by a local axial chord, acircumferential coordinate scaled by the local axial chord and thefirst, second and third span locations, wherein the local axial chordcorresponds to a width of the airfoil between the leading edge and thetrailing edge at respective ones of the first, second and third spanlocations.
 2. The turbine blade according to claim 1, wherein theairfoil is a low pressure turbine blade.
 3. The turbine blade accordingto claim 1, wherein the airfoil is a first-stage, second-stage,third-stage, fourth-stage or a fifth-stage low pressure turbine blade.4. The turbine blade according to claim 1, wherein the airfoil has acold, uncoated airfoil surface at nominal definition.
 5. The turbineblade according to claim 1, further comprising a coating on the airfoil.6. A low pressure turbine blade comprising: a platform; and an airfoilextending in a spanwise direction from the platform to a tip, theairfoil having cross-section airfoil profiles defined by an entirety ofCartesian coordinates set forth in Table
 1. 7. The low pressure turbineblade according to claim 6, wherein the Cartesian coordinates in Table 1have Cartesian coordinate values provided for a cold uncoated conditionat nominal definition.
 8. The low pressure turbine blade according toclaim 6, wherein the tip is shrouded.
 9. The low pressure turbine bladeaccording to claim 6, further comprising a coating over the externalairfoil surface.
 10. A gas turbine engine comprising: a low pressure(LP) turbine configured to drive a low pressure compressor; wherein theLP turbine comprises at least one stage of turbine blades, wherein atleast one of the turbine blades of the at least one stage comprises anairfoil having leading and trailing edges joined by spaced-apartpressure and suction sides extending from a platform in a span directionto a tip; and wherein the airfoil has cross-section profiles defined byan entirety of Cartesian coordinates set forth in Table 1, the Cartesiancoordinates provided by an axial coordinate scaled by a local axialchord, a circumferential coordinate scaled by the local axial chord, anda span location, wherein the local axial chord corresponds to a width ofthe airfoil between the leading and trailing edges at the span location.11. The gas turbine engine according to claim 10, wherein the airfoilhas a cold, uncoated airfoil surface at nominal definition.
 12. The gasturbine engine according to claim 10, wherein the at least one stage ofturbine blades includes between 1 to 5 stages, and wherein at least oneof the 1 to 5 stages comprises the at least one of the turbine blades.13. The gas turbine engine according to claim 12, wherein the at leastone of the turbine blades is a second stage LP turbine blade.
 14. Thegas turbine engine according to claim 10, wherein the at least one stageof turbine blades includes an array of seventy-one (71) turbineairfoils.
 15. The gas turbine engine according to claim 10, wherein thegas turbine engine is a turboprop or a turboshaft engine having 1 to 3independently rotatable spools.